The present invention relates generally to satellite attitude control and, more specifically, to satellite attitude control methods and apparatuses that control the satellite without imparting any significant in-track orbital disturbances. The invention effects attitude control by using electromagnets and thrusters in combination to control the orientation of low earth orbit satellites with respect to their pitch, roll and yaw axes.
Control of the attitude of a spacecraft, such as a satellite, is necessary so that the antennas, which often have narrow beams, are pointed correctly at Earth. A number of factors tend to cause the spacecraft to deviate from the desired rotational orientation, including atmospheric drag, which is the dominant source of error for Low-Earth Orbiting (LEO) satellites. Solar pressure acting on the antennas, spacecraft body, and solar sails may also create rotational forces. Earth""s magnetic field can produce forces on the satellite if it has elements which are magnetic. Because the satellite moves around Earth""s center in its orbit, the forces described above may vary cyclically through the orbit period, and perhaps over a twenty-four hour period. All of these sources of disturbance torque must be actively counteracted by the on-board control system.
In a three axis stabilized satellite, gas jets or xe2x80x9cthrustersxe2x80x9d are typically used to position the satellite with respect to the pitch, roll and yaw axes. On some satellites, one pair of thrusters is needed for each axis to provide rotation in the pitch, roll, and yaw directions, while velocity increments can be provided by using one thruster to act on a given axis.
FIG. 1 shows the reference Cartesian axes XR, YR, ZR) with a satellite 10 at the origin. The ZR axis is directed toward the center of Earth and is in the plane of the satellite orbit. It is aligned along the local vertical of the sub-satellite point. The XR axis is tangent to the orbital path and lies in the orbital plane. The YR axis is perpendicular to the orbital plane. xe2x80x9cRollxe2x80x9d is defined as rotation about the XR axis (xcfx86); xe2x80x9cpitchxe2x80x9d is rotation about the YR axis (xcex8), and xe2x80x9cyawxe2x80x9d is rotation about the ZR axis ("psgr").
The satellite 10 must be stabilized with respect to the reference axes to maintain accurate pointing of its antenna beams. The axes XR, YR, and ZR are defined with respect to the location of the satellite 10, while the second set of Cartesian axes shown in the FIG. 1, X, Y, and Z, define the orientation of the satellite. Changes in satellite attitude cause the angles xcfx86, xcex8, and "psgr" to vary as the X, Y, and Z axes move relative to the fixed reference axes XR, YR, and ZR. The Z axis is usually directed toward a reference point on Earth, called the Z-axis intercept. The location of the Z-axis intercept defines the pointing of the satellite antennas. The Z-axis intercept point may be moved to re-point all the antenna beams by changing the attitude of the satellite with an attitude control system.
There are several ways to make a spacecraft or satellite stable when it is in orbit. For satellites known as xe2x80x9cspinners,xe2x80x9d the entire body of the satellite is rotated at 30-100 rpms to provide a gyroscopic action, which in turn maintains the spin axis in the same direction. In another attitude control method, three or more xe2x80x9cmomentum wheelsxe2x80x9d are mounted on the spacecraft body in order to provide a direct source of control torque. Each momentum wheel consists of a solid disk driven by a motor to rotate at high speed within a sealed, evacuated housing. When control torque is necessary to correct the satellite attitude, the opposite torque is applied to the momentum wheel assembly. With three momentum wheels, rotation of the satellite about each axis can be commanded from Earth by increasing or decreasing the appropriate momentum wheel speed. Since external torque disturbances will tend to cause the speed of the momentum wheels to increase over time, this method requires the use of another source of torque such as thrusters or magnetic torque rods in order to keep the momentum wheels from an overspeed condition.
A momentum bias attitude control system employing a single momentum wheel may include infrared sensors on the spacecraft that are aimed at a reference, such as the outer edge of Earth""s disk, to provide a reference attitude. The reference attitude is compared to a desired attitude. Any error signal generated by this comparison is processed to derive control signals for the motors that drive the momentum wheel for pitch control, and magnetic torque rods for roll and yaw control. Such systems are generally limited in their ability to maneuver about the roll and yaw axes.
Magnetic xe2x80x9ctorque rods,xe2x80x9d in the form of elongated electromagnets, have long been used as part of the momentum management systems of momentum bias controllers for satellites, but only to dampen the wobble generated by the rotating momentum wheels.
One significant advantage of a momentum bias attitude control system is that the combination of torque rods and momentum wheels for control does not impart disturbances to the satellite orbit motion as would thrusters. One disadvantage is that since such wheels must be kept spinning throughout the life of the satellite, they must be designed very carefully and conservatively in order to avoid reliability problems. In the event of a momentum wheel failure, satellites would have to use their thrusters to make attitude adjustments. However, this would be undesirable due to the rapid consumption, and eventual premature exhaustion of propellant as well as undesirable disturbances to the satellite orbital motion.
Examples of attitude control systems can be found in several United States patents. In U.S. Pat. No. 4,521,855 to Lehner et al., a system is disclosed that estimates yaw error, and roll and yaw disturbance torques from measured roll error and yaw momentum on a continuous on-orbit basis. the information can be used to continuously correct for yaw error by activating a xe2x80x9cmagnetic torquer.xe2x80x9d Two control loops are employed. In a xe2x80x9cfastxe2x80x9d loop, nutations are damped by changing momentum wheel speed, while in a xe2x80x9cslowxe2x80x9d loop, yaw error corrections are made by adjusting current to the magnetic torquer.
U.S. Pat. No. 5,308,024 to Stetson describes a satellite attitude control system that employs pitch, roll, and yaw momentum wheels. The system provides attitude control in the event of a loss of the inertial yaw attitude reference, based on modeling of the roll-yaw rigid body dynamics and the roll-yaw orbit kinematics. This system is of particular use when the Earth sensor assembly (ESA), which normally provides yaw reference, fails. U.S. Pat. No. 5,610,820 to Shankar et al. describes a zero momentum attitude control system in which the torque required about a control axis to maintain the desired attitude is determined, and then the thrusters and magnetic torquers are activated to maintain the desired attitude.
A continuing need exists for improved attitude control systems that can effect attitude control without use of momentum wheels or other moving parts, and in particular, for low Earth orbit satellites in a polar orbit. Systems which improve propellant consumption efficiency are also in need.
An object of the present invention is to provide a method and apparatus for controlling the attitude of a satellite or other spacecraft without imparting any significant in-track orbital disturbances.
Another object of the present invention is to provide a satellite attitude control method and apparatus which provide improved propellant consumption performance.
Still another object of the present invention is to provide a satellite attitude control method and apparatus in which no spinning parts are required to control the satellite pitch attitude.
These objectives are met by providing a method of controlling attitude of a satellite which includes the steps of determining whether the measured pitch attitude deviates from a commanded pitch attitude, and powering a torque rod in proportion to the pitch attitude deviation, and in conjunction with Earth""s measured or predicted magnetic field, to produce a magnetic dipole moment which tends to correct the satellite""s attitude.
A separate, simultaneously operating control includes the steps of determining whether the measured roll and yaw attitude rates (time derivative of the roll and yaw attitude measurements, respectively) deviates from the commanded equivalents, and powering a separate torque rod in proportion to the rate error. Again, the magnetic dipole moment generated by the torque rod interacts with the Earth""s magnetic field to generate a correcting torque on the satellite.
Both of these controllers operate in conjunction with thruster-actuated control methods. The cyclical nature of Earth""s magnetic field makes the availability of control torque in a given direction also vary cyclically. In order to make certain that the satellite remains in control even when external disturbances overpower the torque rods, thruster activity remains an optional control method.
Other objects and features of the invention will become more apparent from the following detailed description when taken in conjunction with the illustrative embodiments in the accompanying drawings.